Modern rocket propulsion systems can be classified according to the type of energy source: chemical, nuclear, and solar. Chemical rocket propulsion uses the energy from a high-pressure combustion reaction of propellant chemicals, which heats reaction product gases to very high temperatures. These gases are then expanded in a nozzle and accelerated to very high velocities, which, in turn, bring rockets to high velocities in an opposite direction. Nuclear propulsion, using a fission reactor, a fusion reactor, or directed radioactive isotope decay, has been investigated but remains largely undeveloped. Solar propulsion may use solar panels to heat a gas. The expanded gas can be expelled through an exhaust nozzle, as with chemical propulsion.
Chemical propulsion techniques are typically divided among those using liquid propellants and those using solid propellants. Gaseous propellants and hybrid propellant systems also exist. Typically, liquid propellant rocket engines feed a propellant under pressure from tanks into a combustion chamber. Solid propellant engines, in contrast, store a propellant “grain” in the combustion chamber, the exposed surface of which burns smoothly at a predetermined rate. Combustion chamber conditions therefore vary with propellant type. The techniques applied to control thrust of the various types of rocket engines historically vary to accommodate for the different mechanics of liquid versus solid propellants. Methods for optimizing nozzle efficiency are more developed in the field of liquid propellant engines than in solid propellant motors.
Methods for initiating and stopping liquid propellant rocket engines and for varying the thrust of these liquid engines during operation and flight are described in U.S. Pat. No. 3,897,008; granted Jul. 29, 1975, to Donald G. Nyberg and Ronald F. Dettling entitled “Liquid Fuel Injector System” which is hereby incorporated by reference in its entirety.
Systems providing improved efficiency for liquid rocket engines using expansion-deflection (ED) nozzles and plug nozzles are described in Huzel, Dieter K. and Huang, David H., Design of Liquid Propellant Rocket Engines. Washington D.C.: NASA Science and Technical information Office, 1967, pp. 89-95. The plug nozzle replaces a traditional nozzle exit cone with a spike centerbody. Exiting gases pass through a throat, and then travel down the surface of the spike to converge in a direction opposite that of rocket trajectory.
The use of an ED nozzle is elaborated in Sutton, George P.; Rocket Propulsion Elements, 6th Edition, John Wiley and Sons (1992). As stated therein, “[t]his behavior is desirable at low altitudes because the atmospheric pressure is high and may be greater than the pressure of the exhaust gases. When this occurs, the exhaust is forced inward and no longer exerts force on the nozzle walls, so thrust is decreased and the rocket becomes less efficient. The centerbody, however, increases the pressure of the exhaust gases by squeezing the gases into a smaller area thereby virtually eliminating any loss in thrust at low altitude.”
Liquid propellant engines have improved performance over a wide range of pressure ratios using systems such as those described in Sutton and Huzel and Huang. A recent improvement is described in U.S. Pat. No. 6,591,603 B2, granted Mar. 13, 2003 to Gordon A. Dressier, Thomas J. Mueller, and Scott J. Rotenberger, entitled “Pintle Injector Rocket With Expansion-Deflection Nozzle” (hereinafter “Dressler”). Dressler describes a liquid rocket engine with a variable thrust injector and an ED nozzle to improve performance. In the Dressler system, a throat is formed at one end of a combustion chamber through which hot gases escape. A rod runs through the throat, and a deflector is formed at the end of the rod, downstream of the throat. A nozzle exit cone extends from the throat. Thus, exiting gases pass through the throat and are deflected by the deflector. The deflected gases then pass along the walls of the nozzle exit cone, which direct them in a direction opposite the trajectory of the rocket.
While systems such as the above have improved liquid engine rocketry, no liquid rocket engine design has adequately leveraged improved techniques to provide a simple and powerful engine with both high efficiency over a wide range of backpressures and easily controlled thrust. Such an efficient and versatile rocket engine would provide significant gains in many rocketry applications.
Techniques such as those described above are less developed in the field of solid propellant rocket motors. Designs for use in future generation Army tactical missiles have been investigated and tested, as reported in Burroughs, Susan L. et al, “Pintle Motor Challenges for Tactical Missiles”, AIAA Paper 2000-3310, July 2000. These designs use a pintle that extends into the throat or just upstream of the throat of a conical expansion nozzle. The pintle is attached to a control system that can move the pintle forwards and backwards within the combustion chamber, thereby varying the throat area. The size of the throat area is related to chamber pressure and thrust of the solid rocket motor. After passing through the variable throat area, the exhaust gases are expanded in a conventional nozzle (e.g., conical, bell, Rao, etc) to produce thrust against the walls of the nozzle. A “nozzle pressure ratio” commonly used to characterize the conditions under which a rocket operates is the ratio of internal chamber pressure to external (ambient) pressure against which the rocket exhausts.
Conventional rocket nozzles must be designed to optimize nozzle efficiency at a given nozzle pressure ratio. Nozzle performance (i.e., the efficiency with which a nozzle converts thermal energy of the heated gases in the chamber into thrust-producing, directed kinetic energy of the exhausted gases) typically degrades at nozzle pressure ratios other than the “design,” or optimal, pressure ratio.
As an example, consider a rocket with a constant chamber pressure, a fixed throat area and a conical nozzle which is used to launch a payload through the earth's atmosphere. As the rocket ascends, the ambient pressure into which the motor exhausts (atmospheric pressure) will decrease, thus increasing the nozzle pressure ratio. Nozzle efficiencies at pressure ratios other than the design ratio will be lower than optimal, so rocket designers must choose the pressure ratio “design point” to give the best average performance over the range of expected pressure ratios.
A class of nozzles called “plug” nozzles or “aerospikes,” with a fixed-position centerbody, or spike, that extends downstream of the combustion chamber throat, have the characteristic that nozzle efficiency remains relatively high as a rocket motor with a constant chamber pressure moves through varying ambient pressure conditions. These nozzles are therefore known as “altitude compensating” nozzles.
Nozzles with moveable pintles affect nozzle pressure ratio in a different way, but suffer nonetheless from loss of nozzle efficiency at “off-design” pressure ratios. In this class of nozzles, the pintle is used to vary the throat area, and thus the thrust of solid propellant motors. In varying throat area, these nozzles vary the chamber pressure, and thus the propellant burn rate, with the ultimate effect of varying thrust. However, because the pintle is used in combination with a cone nozzle, varying pressure ratios force rockets of such a design to operate at sub-optimal pressure ratios. Thus thrust control, or “throttling” is achieved at the cost of nozzle efficiency.
Thus, theory and test results demonstrate that the tested designs cannot maintain high performance over a wide range of nozzle pressure ratios. This is largely because such designs suffer from efficiency losses due to expansion problems in a fixed nozzle exit cone or bell nozzle configuration. Regardless of whether the change in nozzle pressure ratio occurs because of decreasing exhaust pressure (increasing altitude) or decreasing chamber pressure (thrust throttling), nozzle efficiency suffers due to non-optimal nozzle expansion at off-design nozzle pressure ratios. Performance losses of up to 30% off of optimal efficiency can occur at off-nominal pressure ratios. To date, no method has been identified for maintaining near-optimal nozzle efficiency while varying thrust over a wide range.
In summary, both liquid and solid rocket motor designs have failed to realize their full potential in providing both high efficiency over a wide range of pressure ratios, and thrust control. Such an efficient and versatile solid, liquid, or other propellant type rocket would provide significant gains for rockets used for commercial and military spacecraft launches, as well as missile launches used for both conventional and anti-terrorism warfare.